In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot gases are then channelled towards a gas turbine which transforms the energy from the hot gases into work for powering the compressor and other devices which converts power, for example an upstream fan in a typical aircraft turbofan engine application, or a generator in power generation application.
The gas turbine stages include stationary vanes which channel the combustion gases into a corresponding row of rotor blades extending radially outwardly from a supporting rotor disk. Each rotor blade typically comprises a lower root, for coupling the blade with the turbine rotor, and a hollow aerofoil exposed to the combustion gases channelled through the vanes. Aerofoils may be designed and manufactured hollow in order to save weight, to change its eigenfrequency or to include a cooling circuit therein. In the latter case, the cooling gas which circulates inside the cooling circuits is typically bleed air from the compressor discharge. Between the root and the hollow aerofoil, each rotor blade typically further comprises a platform, having an upper surface from which the aerofoil extends. The upper surface of the platform is therefore also exposed to the hot combustion gases channelled through the vanes.
Blade platforms at the first stages of a gas turbine are in contact with combustion gases having high temperature, typically higher than 900° C. Usually the blade platform is covered by a thermal barrier coating (TBC) for protecting the blade platform from corrosion. As long as the temperature in the bond coat under the TBC layer is under 920° C., then the TBC layer is stable and blade can be effectively protected from corrosion. However, the need for improvement in turbine efficiency and output power makes the gas temperature at the turbine inlet hotter and hotter. Therefore, known solutions, as providing a TBC layer for coating the platform, cannot be considered any more reliable for preventing corrosion of the blades during all the turbine lifecycle.
Other solutions for the cooling of blade platforms are shown in U.S. Pat. No. 4,672,727, GB 2 244 520 and EP 1 574 670, but cannot yet considered optimal. In U.S. Pat. No. 4,672,727 the platform is could by means of the air trapped in compartments formed between the undersides of the platforms and the rim of the disk. In GB 2 244 520 a design is disclosed for cooling a limited portion of the platform of a blade. In EP 1 574 670 a portion of the air flow channelled to cool the aerofoil of a blade is diverted through a grove provided in the rotor to cool the platform.
It is desirable to provide a new design for gas turbines where the blade platforms are cooled by new and more efficient cooling techniques.